Facilities

The Variable Mach Number Supersonic Wind Tunnel is capable of providing a test flow with a continuously variable Mach number in the range of 1.3 to 3.1. The facility (see left image) consists of a 300 psi compressed air source, settling chamber, variable nozzle block (see right image detail), test section, and diffuser. Formally a pilot facility for the NASA Langley Supersonic Unitary Plan Wind Tunnel, the facility is used to study supersonic aerodynamics and compressible flow phenomena. It has been used in the past to simulate fuel-air mixing downstream of a pair of swept ramp fuel injectors. The test time of the facility is typically 2.5 minutes and it has a test section cross section of 3 × 3 inches. The facility is equipped with low frequency pressure measurement and shadowgraph flow visualization systems. The wind tunnel is used for both research and undergraduate instructional laboratories.

This unique hypersonic wind tunnel at the Aerospace Research Laboratory is used to produce speeds to Mach 16 continuously.  The facility consists of a large vacuum tank into which nitrogen is introduced through a sonic orifice.  The large pressure ratio across the orifice produces an underexpanded jet.  Models are placed along the centerline of the jet for testing from Mach 1 to 16.  NASA funded research has been done related to spacecraft entering the Martian atmosphere using retropropulsion rockets and reaction control jets.  Current testing employs a model consisting of an array of Mach 5 jets impinging on a flat plate to simulate the impact of spacecraft retrorockets on the surface of the International Space Station (ISS).  Such impacts have been observed during docking maneuvers, causing damage to the ISS solar panels.

Measurements are made in the flow field around these models using a nonintrusive optical technique called Planar Laser-Induced Iodine Fluorescence (PLIF). This technique can measure all thermodynamic parameters as well as velocity.  It is the only technique capable of such quantitative, spatially resolved measurements in these highly rarefied flow fields.

The Inertial Particle Separator (IPS) Wind Tunnel is a two-dimensionally simplified test facility that approximates the particle filtration systems found on helicopter jet engines.  The wind tunnel (on the left in the image) employs a pull-down configuration whereby the flow enters a contraction section followed by the test section with optical access covering a length of 30 cm and spanwise width of 20 cm.  The inlet section is capable of flow inlet speeds of up to 150 m/s and separates the incoming flow into two legs: a scavenge flow (up to 300 SCFM) and a core flow (up to 1500 SCFM).  The two flow legs are independently controlled with a PID automated controller. These flow conditions simulate realistic full-scale engine inlet conditions in terms of flow velocity, length scales, and test particles, which are released above the contraction section as a dilute particle-laden jet (right). A variety of separator geometries can be tested and the available diagnostics for this facility allow particle separation efficiency measurements as well as velocity field descriptions using Particle Image Velocimetry (PIV) and flow visualization using High-Speed Video (20,000 fps) data capture.

The ASOS Shock Tube installation is a large scale shock tube facility with a test section cross sectional area of 200 × 200 mm. This test section is specially engineered to withstand the dynamic loadings at the facility-design envelope with a fully transparent (polycarbonate) construction. The shock tube is supported in a horizontal orientation and employs a double-diaphragm bursting arrangement for controlled operation. Shock Mach numbers up to 3.5 and a flow dynamic pressure of 2 MPa can be generated. The 2 m long driver section ensures that uniform flow conditions exist over a time period of 3 to 8 ms.

The simultaneous use of three high-speed cameras provides a unique opportunity to observe the reaction of the target to the shock using different points of view with frame rates up to 100 MHz. Application of different visualizing techniques allows one to observe the reaction using different methods (photography, shadowgraphy) and to measure density gradient distributions during the process (Schlieren visualization).

The PISTL facility is a shock heated wind tunnel that is capable of operating in both shock tube mode and expansion tube mode. Formerly a federally owned facility located at NASA Langley Research Center, the wind tunnel is being installed and recommissioned at the ARL. The facility is approximately 50 feet long and has an internal diameter of 3.5 inches. It can generate flow velocities exceeding 4000 m/s, (close to 9000 mph), and is therefore well suited to research on the aerodynamics associated with hypervelocity flight, including atmospheric re-entry of a spacecraft and the entry of a spacecraft into the atmospheres of other planets in the solar system. It is also capable of high Mach number, high enthalpy boundary layer experiments and supersonic combustion experiments. The research applications of this facility go beyond hypersonic aerodynamics and combustion to include chemical kinetics, plasmas, and blast wave damage and mitigation, as well as diagnostics and measurement technique development. The facility complements the ASOS shock tube facility at ARL due to the higher flow velocities and enthalpies that can be generated. The facility is being installed with an array of high frequency (MHz) pressure transducers as well as a state-of-the-art, high-speed (MHz) digital data acquisition system. High speed (MHz) Schlieren and shadowgraph flow visualization are also available.